Turbine cooling device



Sept. 13, 1966 D. H. slLvERN ETAL 3,271,952

TURBINE COGLING DEVICE Filed Aug. 14, 1964 5 Sheets-Sheet l F/PST 577966- Sept. 13, 1966 D. H. SILVERN ETAL 3,271,952

TURBINE COOLING DEVICE Filed Aug. 14, 1964 5 Sheets-Sheet 2 164 HZ n; '10

Sept 13 l965 D. H. slLvERN ETAL 3,271,952

TURBINE COOLING DEVICE s sheets-sheet s Filed Aug. 14, 1964 United States Patent G 3,271,952 TURBINE COOLING DEVICE David H. Silver-n, North Hollywood, Calif., and Kenneth E. Nichols, Arvade, Colo., assignors to Sandstrand Corporation, a corporation of Illinois Filed Ang. 14, 1964, Ser. No. 389,622 9 Claims. (Cl. 6ft-39.66)

This invention relates generally to rotary turbines, and more particularly to a cooling device for a rotary turbine.

In turbines which employ a high temperature gaseous propellent to rotate the turbine wheel, the possibility of high temperature failure of the turbine blades and the turbine disk becomes a critical problem. Unless some provision is vmade to maintain the turbine wheel and blade temperatures below the temperature of the hot gaseous propellent, failure of the turbine will result. In many applications of a turbine as the source of motive power, it is impractical to use air as the cooling fluid for the turbine disk and blades. One such example is a torpedo, or other underwater objects employing a turbine as a source of power, where air for cooling the turbine wheel is not readily available. Therefore, the present invention is directed to a device for cooling the turbine disk and turbine blades in installations where the use of conventional cooling fluids is not feasible.

It is therefore a primary object of the present invention to provide a new and improved cooling device for a turbine wheel.

Another object of the present invention is to provide a cooling device for a turbine wheel in which a portion of the combustion products of the propellent are cooled and thereafter directed against the turbine wheel to maintain the wheel at a temperature significantly below the temperature of the hot propellent gases impinging on the turbine blades.

A more specific object of the present invention is to provide a cooling device for a turbine wheel in which the critical metal components are cooled with a gaseous propellent which has been previously cooled in a heat exchanger with a secondary cooling medium. By bleeding off a small portion of the gaseous propellent in the distributor, which conveys the lhot gases from the propellent combustor to the turbine hot gas nozzles, and thereafter cooling the bled propellent and directing it against the critical portions of the turbine wheel an effective and -simple cooling device is provided not heretofore known in the prior art.

A further object of the present invention is to provide a new and improved cooling device for the critical metal components of a turbine wheel in which the turbine is cooled with a portion of the gaseous propellent after the propellent has passed through the turbine blades. When a greater volume flow of cooling fluid is required, the exhaust gases may be cooled in a heat exchanger and directed upon the turbine wheel to maintain the turbine wheel at a relatively constant temperature significantly below the hot gas inlet to the turbine blades thereby preventing failure of the turbine wheel even under maximum power and maximum depth conditions.

Another object of the present invention is to provide -a new and improved stationary shroud surrounding the turbine wheel having -means for directing a cooling fiuid against the turbine blades immediately after the blades have been impinged by the hot combustion gases. The use of a cooling device that directs the cooling iiuid directly against the blades in this fashion serves the twofold purposes of scavenging the residual hot gas from the blade passages and also leaving some relatively cool gas in these passages to absorb heat during the remainder of rotation of the wheel until the blades again return to the hot gas zone. This has been found to be a par- ICC ticularly effective cooling device as it prevents some of the heat input to the wheel which occurs in conventional cooling devices.

A still further object of the present invention is to provide a buffer chamber between the exhaust gases and a cooling shroud on the exhaust side of the turbine wheel so that the gases in the buffer chamber insulate the cooling shroud from the exhaust gases and provide a more efiicient heat ow from the exhaust side o-f the turbine disk.

A more specific object of the present invention is to provide a hot gas turbine having a rotary turbine wheel with blades extending peripherally from a disk, source of hot gaseous propellent, a heat exchanger connected to a source of cooling fiuid, means connecting a portion of the hot gaseous propellent to the heat exchanger to cool a portion of the propellent, a first shroud adjacent one side of the turbine wheel, nozzles in the fir-st shroud for directing the hot gases against the turbine blades, the nozzles extending about a portion of the shroud, cooling fluid guide means in the shroud adjacent the nozzles for directing cooling iiuid against the turbine blades after they pass through the hot gas portion, a cooling passage in the shroud radially inward from the guide mean-s for directing cooling fluid against the turbine disk, means connecting the cooling fiuid g-uide means and the cooling passage to receive the cool gas from the heat exchanger, clearance means in the shroud adjacent the wheel permitting cooling fluid from the passage to iiow radially across the disk and join the cooling iiuid in the guide means before the latter pases through the turbine blades, second shroud means adjacent the exhaust side of the turbine wheel, passage means in the second shroud for directing cool gas against the exhaust side of the turbine disk radially inward from the blades, clearance means in the second shroud means adjacent the turbine wheel permitting cooling fluid from the passage means to flow radially outward to cool the disk and join the cooling fluid from the first shroud means passing through the turbine blades, an annular buffer chamber adjacent the second shroud means and axially spaced therefrom away from the turbine wheel, and conduit means in the second shroud connecting the buffer chamber so that the gases passing through the blades and from the second shroud clearance means flow out through the buffer chamber and shield the second shroud from the hot exhaust gases thereby preventing any significant heat transfer to the exhaust side of the turbine wheel.

Other objects and advantages will become readily apparent from the following detailed description taken in connection with the accompanying drawings, in which:

FIGURE 1 is a partially schematic drawing of a twostage turbine incorporating the cooling device of the present invention;

FIGURES `2 and 3 are schematic representations o-f the hot gas and cooling gas stages on the bearing and exhaust sides, respectively, of the turbine shown in FIG- URE l;

FIGURE 4 is a fragmentary, partially schematic crosssection of the turbine shroud adjacent the cooling gas stage;

FIGURE 5 is a schematic view of a modification of the cooling device in FIGURE l;

FIGURE 6 is a cross-sectional elevation of a single stage turbine showing the cooling shroud adjacent the turbine wheel;

FIGURE 7 is a cross-sectional, developed plan view taken generally along line '7-7 of FIGURE 6 showing the hot gas nozzles and the cooling gas nozzles;

FIGURE 8 is a side elevation partially in cross-section taken generally along line 8 8 of FIGURE 6 showing the inlet shroud of the turbine; and

FIGURE 9 is a side elevation partially in cross-section taken generally along line 9-9 of FIGURE 6 showing the exhaust shroud of the turbine. While illustrative embodiments of the invention as shown in the drawings will be described in detail herein, the invention is susceptible of embodiment in many different forms and it should be understood that the present disclosure is to be considered as an exemplilieation of the principles of the invention and is not intended to limit the invention to the embodiments illustrated. The scope of the invention will be pointed out in the appended claims.

Reerring to FIGURES l to 4 wherein the cooling deviceis shown -schematically with a two-stage turbine, a turbine wheel 10 is adapted to rotate about an axis of rotation 11. The turbine wheel 1t) consists of a disk 12 having radially extending blades 13 extending 360 about the annular periphery of the disk 12.

A propellent tank 15 is adapted to contain a gaseous propellent supplied to a combustor 17 by pump 18 driven through gear box 19 by the turbine wheel 10 through a suitable driving connection. A turbine wheel controller 26 maintains the desired speed of rotation of the turbine 1t). The partially dotted lines in FIGURE l indicate the flow of propellent, such as line 22 connecting the propellent pump 18 to the combustor 17. The combustor 17 burns the propellent from tank 1S and delivers the hot decomposition products to the turbine blades 13 through distributor 25 connecting the combustor 17 to an inlet shroud 26 adjacent the turbine wheel 10. Hot combustion gases from the distributor 25, indicated by line 28 in the distributor, pass through nozzles, not shown in FIGURE 1, in shroud 26 which dire-ct the hot gases against the turbine blades 13 over a portion of the shroud periphery designated by the numeral 32 as the first stage inlet in FIGURE 3. As the turbine in FIGURES l to 4 is a two-stage turbine, the exhaust 33 in FIGURE 2 from the first stage is turned by suitable ducting, not shown, and readrnitted to the turbine blades 13 through nozzles 37 .constituting the second stage inlet 38 shown in FIG- URE 2. The gases from nozzles 37 after passing through the blades 13 are exhausted from the turbine at the exhaust side of the turbine Wheel 10. The exhaust 4t), shown in FIGURE 3, from the second stage requires a greater annular portion of turbine periphery because of the increased volume of the gases exhausted from the iirst stage 33.

Referring to FIGURES 2 and 3 as the hot-gas stages of the turbine require only a portion of the turbine annulus for their inlets and exhaust, this is a partial admission turbine.

As shown in FIGURE l, a portion of the hot decombustion products from the combustor 17 .are bled off from the distributor 25 through line 41, the dotted lines indicating the flow of coolant in FIGURE l. Suitable conduits, not shown, may be provided to transfer the hot gas from the distributor. The hot gas in line 41, which by way of example may be about 210 pounds per hour, flows through `a heat exchanger 45 Where the hot combustion gases are cooled and then delivered to the turbine 10 through line 46.

A secondary lcooling fluid, such as water, is delivered to the heat exchanger 45 to cool the hot combustion gases. The use of water as a secondary cooling uid is particularly advantageous when the turbine is in an underwater installation such as a torpedo where a source of sea Water is readily available. A sea water pump 48 delivers water through line 49 to suit-able heat exchange tubes in the heat exchanger 45 and also delivers water to the combustor 17 through line 50. The barbed lines indicate the flow of cooling water in the turbine in FIGURE l. The distributor 25 has a water jacket passage 52 which conveys the cool water to the shroud 26 adjacent the turbine Wheel 10 to maintain a compatible shroud temperature.

The cooled combustion gases flowing in line 46 constitute the cooling iluid for the turbine and are delivered to ducts 55 and 56 adjacent the bearing side of the turbine Wheel and duct 58 adjacent the exhaust side of the turbine wheel 1li. An exemplary division of flow between the ducts is; l5 pounds per hour for duct 55, 180 pounds per hour for duct 56, and l5 pounds per hour for duct 58.

Referring to FIGURES 1 and 4, ycooling fluid in duct 55 passes in an annular clearance spa-ce 61 between the shroud 60 and the disk 12 and iloWs radially across the surface 63 of disk 12 cooling this critical portion of the turbine wheel 10. The greater quantity of cooling fluid ow from duct 56 directly to the turbine blades, rather than across the disk 12, as shown more clearly in FIG- URE 4 by arrow 65 which indicates the flow from duct 56. The flow from duct 56 joins the radial flow in clearance space 61 before passing through the turbine blades 13. Suitable nozzles may be provided in the shroud 60 between the first stage exhaust 33 and the second stage inlet .38, defining a blade cooling gas inlet 68, as shown in FIGURE 2. Cooling fluid nozzles may also be provided between the second stage inlet 3S and the iirst stage exhaust 33, dening another cooling gas inlet 69. Therefore, while the cool iluid from duct 55 passes radially in clearance space 61 all around the disk 12, cooling lluid from duct 56 is injected in the blades 13 only over a portion of the periphery of the wheel 10. The cool gas flowing across the blades from the cooling gas inlets 68 and 69 serve to scavenge the hot gases between the blades immediately after passing through the hot gas stages of the turbine, and also leave cool air between the blades to maintain a low blade temperature before passing through the next hot gas stage.

The -cooling fluid exits from the blades in cooling fluid exhaust portions 71 and 72, as shown in the stage diagram in FIGURE 3. A clearance 75 in shroud 26 directs the flow of cooling iluid from the blades radially inward along the exhaust side of the turbine wheel 10.

Shroud 26 has an annular flange 77, as shown in FIG- URE l, which forms an annular buifer chamber 78 between the hot gas exhaust and flange 80 on shroud 26. Conduits S1 are formed in iiange 80 permitting the flow of cooling gas from the turbine blades and the ow of cooling gas from duct 58 yto pass into the buffer chamber 78 and from there to exhaust. In this manner the bulfer chamber 78 serves to insulate or shield the shroud flange 80 and clearance '75 adjacent the duct 58 from hot exhaust gases leaving the turbine, and thereby improve the thermal efficiency of the radial gas ow along the exhaust side of the disk 12 and help to maintain a lower disk temperature. It should be understood that FIGURE 4 is partially schematic and therefore does not show the exact construction ofthe shrouding 26 shown in FIG- URE 1.

In FIGURES 1 to 4, hot combustion gases are drawn off the distributor 25 before passing through the turbine blades for the purpose of providing cooling fluid for the turbine wheel. In some instances Where a greater volume of cooling iluid is required it is more eiiicient to use the exhaust gases, i.e. the hot combustion gases after passing through the turbine blades, as the cooling fluid. Viewing FIGURE 5, wherein a single stage turbine is generally indicated by the numeral 110, hot combustion gases supplied by a propellant tank 111 and a combustor 112 pass through hot gas nozzles 113 which direct the gases against the turbine blades 115 on turbine wheel 116. The gases then are exhausted rather than reentered in another section of the turbine wheel 116. A portion of the hot exhaust gases are drawn off by suitable ducting, not shown, through line 118 to a heat exchanger 119 where they are cooled by sea water or another suitable medium, as discussed above with respect to FIGURES l to 4. The cooled gas from the heat exchanger is supplied to the critical portions of the turbine wheel 116 through lines 121 and 122 in the same fashion as discussed above with respect to FIGURE 1.

Referring to FIGURES 6 to 9, the inlet and exhaust shrouding of the present invention is shown in detail in a single stage turbine generally indicated by the numerial 150. Turbine wheel 151 has a disk portion 152 and radially extending blades 153 about the periphery thereof. The turbine wheel 151 is mounted for rotation in suitable bearings, not shown, seated within support member 155. Surface 157 is on the inlet side of disk 152 and surface 158 is on the exhaust side thereof. Shroud members 160 and 161 house t-he turbine wheel 151 on the inlet and exhaust sides thereof respectively. The shroud 160 has a generally annular shape and is fixed to the support member 155 by suitable studs 162. Annular channels 163 and 164 are formed in surface 165 of shroud 160 and communicate with an outlet 166. Viewing FIGURES 6 and 8, an annular fiat plate 170 is fastened to the shroud surface 165 to define annular water cooling passages with the channels 163 and 164 so that the shroud 160 is maintained at a relatively cool temperature.

Referring to FIGURES 6 and '7, -hot gas nozzles 175, 176 and 177 are formed in the inlet shroud 160 over a small angular portion of the shroud thereby defining the turbine 150 as a partial admisison type. The hot gas inlet port 178 directs the hot gases into the inlet portions of the nozzles 175, 176 and 177, and is located in a cut-out portion 180 of the plate 170 as shown in FIGURE 7. A hot gas manifold 180 fixed to port member 178 and plate 170 conveys the flow of hot gases from a source of gaseous propellant, such as that shown in FIGURE l to the hot gas nozzles 175, 176 and 177. The manifold 180 is cooled by a water cooling jacket 181 which also supplies cooling water to the channels 163 and 164, which water exhausts through outlet 166 in shroud 160. Referring to FIGURES 6 and 7, an oblong exhaust port 185 in shroud member 161 on the exhaust side of the turbine wheel 151 directs hot exhaust gases from the turbine blades into exhaust chamber 186 formed by a housing member 187 fixed to shroud member 161 on the exhaust side of the turbine.

Adjacent the hot gas nozzles 175, 176 and 177 in the direction of rotation of the turbine wheel 151, are cooling gas nozzles 190, 191, 192 and 193 fixed in the inlet shroud member 160 and protruding therefrom on the inlet side of the turbine through apertures 195 in plate 170. Cooling fluid ows into nozzles 190 to 193 from the heat exchanger, such as that shown in FIGURES 1 to 4. The cooling fluid from the cooling nozzles scavenges hot gas from the turbine blades 153 immediately after they pass over the hot gas nozzle portion of the shroud 160 and also cools the blades 153 maintaining a blade temperature considerably below the temperature of the hot gases entering the nozzles 175 to 177. For example, this method of cooling for hot gas inlet temperatures of 2500" F. can maintain the blade temperatures below 1400 even under extreme conditions. A portion of the cooling uid after passing through the turbine blades is directed into the exhaust chamber 186 through cooling port 198 in exhaust shroud 161.

Cooling fluid is lalso directed against the disk portion 152 of the turbine wheel 151 to cool the inlet surface 157 thereof. This is achieved by an inlet port 199 in support member 155 adapted to receive a fitting connected to the heat exchanger shown in FIGURE 1, and passages 200 connecting the inlet 199 to an annular passage 201 in stationary shroud ring 202. Passages 203 in shroud ring 202 in annular array therein deliver cooling fluid against the inner portion of the surface 157 which because of the rotation of the wheel 151 flows radially outward in a clearance space 205 between the shroud members 160 and 202 and the surface 157 of the turbine wheel to cool the disk portion of the wheel 151.

Now the uid flowing radially across the disk 152 on the inlet side of the turbine wheel 151 and the cooling fluid from nozzles 190 to 193 pass through and around the turbine blades 153 to the exhaust side of the turbine Wheel 151 and flow radially inward a short distance in clearance space 210 between the shroud 161 and the surface 158 on disk 152. Shroud 161 has a conical web 211 which forms the clearance space 210 and terminates in a central boss 212 concentric with the axis of rotation ofthe turbine wheel 151.

Cooling gas from the heat exchanger also is supplied to fitting 214 connected to a conduit 215 which delivers cooling fluid to the center of the exhaust side of turbine wheel 151 through passage 217 in the boss 212. The cooling fluid from passage 217 flows radially across surface 158 on the disk 152 cooling the same and joins the cooling fluid which has passed through the blades from the inlet side of the turbine wheel 151.

The conical plate 220 forms an annular buffer chamber 221 separating the exhaust chamber 186 from the shroud web 211 and clearance space -210. The joined cooling fluid flows through conduits 222 annularly arranged in web 211 of shroud 161 and passes radially inward in buffer chamber 221 and into exhaust pipe 225 fixed to the plate 220. The buffer chamber 221 serves to insulate or shield the web 211 and clearance space 210 from the hot exhaust gases in chamber 186 so that the cooling fluid flowing radially outward in the clearance space 216 can more effectively reduce the temperature of surface 158 on the exhaust side of turbine wheel 151.

Pipes 227, 228 and 229 are adapted to receive thermometers for measuring the temperatures at various positions adjacent the turbine wheel 151. Shroud 161 is also water cooled by water flowing in annular passage 231 from water inlet fitting 232 adapted to be connected to the sea water pump in FIGURE 1. The water is exhausted from shroud 161 around the exhaust pont 185 as shown in FIGURE 6.

The cooling device maintains turbine Wheel temperatures below that of any cooling device heretofore known and results in variations in blade surface temperature during one complete revolution of less than 5 F. The fluid flowing in the buffer chamber of the exhaust side of the turbine wheel is particularly effective as a heat shield due to the low relative velocity thereof and the consequent low film coefficient.

We claim:

1. In a hot gas turbine having a source of hot gaseous propellant and a turbine wheel having blades extending peripherally from a disk, the improvement comprising; shroud means adjacent one side of the turbine wheel, hot gas nozzle means extending over a portion of said shroud means adjacent the turbine lblades for directing hot gas against the blades, first cooling fluid means in said shroud means extending over a second portion thereof adjacent said blades for directing cooling fluid against said blades, second shroud means adjacent the exhaust side of the turbine wheel and having passage -means therein adjacent the disk, clearance means in said second shroud means adjacent the turbine wheel connecting said passage means and the turbine blades so that the cooling fluid from the b-lades joins the cooling fluid from said passage means, an annular buffer chamber in said second shroud means spaced axially from said clearance means thereby separating the clearance means from the hot exhaust, and passage means connecting said clearance means with said buffer chamber, said first passage means and said chamber being substantially parallel, whereby the joined cooling fluids pass out through the buffer chamber and shield the c-learance means from the hot exhaust gases.

2. In a hot gas turbine having a source of gaseous propellant and a turbine wheel with blades peripherally extending from a disk, the improvement comprising; a first shroud adjacent one side of the turbine wheel, hot gas nozzles in said shroud adjacent said blades and extending over a portion of the periphery of said shroud, cooling fluid nozzles in said shroud adjacent said blades and extending over a second pontion of said periphery, said second portion being adjacent said iirst portion in the direction of rotation of said turbine wheel, passage means in said first shroud adjacent the disk of the turbine Wheel for cooling the disk, clearance means in said iirst shroud connecting said passage means and said irst cooling means so that the co-oling iluid from each joins before passing through the turbine blades, a second shroud adj-acent the exhaust side of the turbine wheel, a cooling iiuid passage in said second shrou-d adjacent the turbine disk, clearance means in said second shroud permitting cooling iiuid from said passage to ow radially across 'the turbine disk, a buffer chamber against said second shroud and spaced axially therefrom away from said turbine wheel, said cooling uid passage in the second shroud means being substantially parallel with said butter chamber, and conduit means connecting said second shroud clearance means with said butter chamber, so that the cooling iiuid from said cooling passage and trom -t-he 'blades passes into the buffer chamber thereby shielding the second shroud from the hot exhaust gases.

3. A hot gas turbine, comprising; a rotary turbine Wheel having blades extending peripherally from a disk, a source of hot gaseous propellant, a heat exchanger connected to a source of co-oling fluid, means connecting a portion of said hot gaseous propellant to said heat exchanger to cool said propellant, first shroud means adjacent one side of thel turbine wheel, nozzles in said shroud means for directing the hot gases against the turbine blades, said nozzles extending about a portion of said shroud, cooling iiuid guide means in said shroud adjacent said nozzles for directing cooling iuid against the turbine blades after they pass through the hot gas portion, a cooling passage in said shroud radially inward from said guide means for directing cooling fluid against the turbine disk, means connecting said cooling fluid guide means and said cooling passage to receive cool gas from said heat exchanger, clearance means in said shroud adjacent the wheel permitting cooling uid from said passage to flo-w radially across the disk and join the cooling fluid in said guide means before the latter passes `through the turbine blades, second s-hroud means adjacent the exhaust side of the turbine wheel, passage means -in said second shroud for directing cool gas against the exhaust side of the turbine disk radially inward from the blades, clearance means in said second shroud means adjacent the turbine wheel permitting cooling gas from said passage means to liow radially outward to cool the disk and join the cooling yiluid from said irst shroud means passing through the turbine blades, an annular buffer chamber adjacent said second shroud means and axially spaced therefrom away from the tur-bine Wheel, and conduit means in said second shroud communicating ywith said buffer cham-ber so that the gases passing through the blades and from the second shroud clearance means flow out through the butter chamber and insulate the second shroud from the hot exhaust gases thereby preventing any signiiicant heat transfer to the exhaust side of the turbine Wheel.

4. In a hot gas turbine having a source of hot gaseous propellant and a turbine wheel'having blades extending peripheral'ly from a disk, the improvement comprising: shroud means adjacent one side of the turbine wheel, |hot gas nozzle means extending over a portion of said shroud means adjacent the turbine blades for directing hot gas against the blades, a second shroud means adj-acent an exhaust side of the turbine wheel and having passage means therein adjacent the disk, means for supplying cooling fluid to said passage means so that the cooling fluid ows radially across the disk, a generally annular, radially extending buffer chamber adjacent Said shroud means spaced axially therefrom for separating said shroud means from the hot exhaust, and passage means connecting said clearance means with said buffer chamber to direct the cooling fluid iiowing radially in one direction across said disk to ow generally radially in the other direction Within said buffer chamber, said first passage means and said chamber being substantially parallel` 5. In a hot gas tunbine for delivering power in an environment Where a source of cooiing air is not readily available, the combination comprising: a self contained self supporting combustion system including a combustor `for burning a propellent and producing hot combustion products Iwithout employing atmosperic air, a turbine wheel mounted for rotation, a distributor connecting said combustor to said turbine wheel for rotation thereof, shroud means adjacent said turbine wheel having passages therein for directing coo-ling tluid adjacent said wheel, a heat exchanger, means for delivering only combustion products to said heat exchanger for cooling said combustion gases, means connecting said heat exchanger to said passages to deliver cooled combustion gases directly to said passages, means for exhausting at least a major portion of the cooling gas from the turbine, and a fluid pump for delivering uid to said heat exchanger to cool said combustion gases, whereby cooled combustion gases provide a cooling iiuid in t-he absence of atmospheric air.

6. A hot gas tur-bine as defined in claim 5 wherein said iiuid pump is a water pump..

7. A hot gas turbine as defined in claim 5 wherein said means for delivering only combustion products to sai-d heat exc-hanger comprises conduit means connected to and receiving gases from said distributor.

l8. A hot gas turbine as deiined in claim 5 wherein said means for delivering only combustion gases to said `heat exchanger comprises conduit means for receiving combustion gases exhausting from said turbine wheel.

9. A hot gas turbine as deiined in claim 5 wherein said shroud means includes an annular porti-on adjacent one side of the turbine Wheel extending from the periphery of said wheel toward the center of said wheel, hot gas nozzles extending over an arcuate portion of said shroud means adjacent said blades for directing hot gases V against said blades thereby delining a partial admission tunbine, said cooling passages including cooling iluid nozzles extending over a second arcuate portion of said shroud means adjacent said blades for cooling the blades, chamber means in said shroud means radially inwardly of said nozzles for directing cooling iiuid against the disk portion of the Wheel, and clearance means in said shroud means for conveying cooling uid radially outward from said chamber means, said clearance means being connected for combining the cooling uid from the chamber with the cooling fluid from the cooling nozzles so that both pass through the turbine blades.

References Cited by the Examiner UNITED STATES PATENTS 1,201,545 ll0/1916 Bischof 60-3952 X 2,354,698 8/1944 Norris 6039-5 2,401,826 6/ 1946 Halford. 2,419,689 4/ 1947 McClintock 60-39.43 2,586,025 2/2952` Godfrey. 2,603,948 7/195'2 vMirns 6ft-39.52 `3,085,396 4/1963 Kent 601-3966 X 3,134,228 5/1964 Wolansky Gli- 39.46 X

MARK NEWMAN, Primary Examiner.

RALPH D. Asrfnam Examiner. 

5. IN A HOT GAS TURBINE FOR DELIVERING POWER IN AN ENVIRONMENT WHERE A SOURCE OF COOLING AIR IS NOT READILY AVAILABLE, THE COMBINATION COMPRISING: A SELF CONTAINED SELF SUPPORTING COMBUSTION SYSTEM INCLUDING A COMBUSTOR FOR BURNING A PROPELLENT AND PRODUCING HOT COMBUSTION PRODUCTS WITHOUT EMPLOYING ATMOSPHERIC AIR, A TURBINE WHEEL MOUNTED FOR ROTATION, A DISTRIBUTOR CONNECTING SAID COMBUSTOR TO SAID TURBINE WHEEL FOR ROTATION THEREOF, SHROUD MEANS ADJACENT SAID TURBINE WHEEL HAVING PASSAGES THEREIN FOR DIRECTING COOLING FLUID ADJACENT SAID WHEEL, A HEAT EXCHANGER, MEANS FOR DELIVERING ONLY COMBUSTION PRODUCTS TO SAID HEAT EXCHANGER FOR COOLING SAID COMBUSTION GASES, MEANS CONNECTING SAID HEAT EXCHANGER TO SAID PASSAGES TO DELIVER COOLED COMBUSTION GASES DIRECTLY TO SAID PASSAGES, MEANS FOR EXHAUSTING AT LEAST A MAJOR PORTION OF THE COOLING GAS FROM THE TURBINE, AND A FLUID PUMP FOR DELIVERING FLUID TO SAID HEAT EXCHANGER TO COOL SAID COMBUSTION GASES, WHEREBY COOLED COMBUSTION GASES PROVIDE A COOLING FLUID IN THE ABSENCE OF ATMOSPHERIC AIR. 